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Sentinel-2A/B Thermal Design - Lessons Learnt from TBTV, LEOP and IOC

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Sentinel-2A/B Thermal Design - Lessons Learnt from TBTV, LEOP and IOC 48th International Conference on Environmental Systems ICES-2018-148 8-12 July 2018, Albuquerque, New Mexico Sentinel-2A/B Thermal Design - Lessons Learnt from TBTV, LEOP and IOC Nadine Buhl1, Martin Altenburg2, Markus Manns3 Airbus Defence and Space GmbH, Claude-Dornier-Straße, 88090 Immenstaad , Germany The Sentinel-2 (S-2) satellite is developed by ESA and forms part of the EU Copernicus program. It carries an innovative wide swath high-resolution Multispectral Imager for a new perspective of our land and vegetation. The combination of high resolution, novel spectral capabilities, a swath width of 290km and frequent revisit times will provide unprecedented views of Earth. The mission is based on a constellation of two identical satellites in the same orbit, 180° apart for optimal coverage and data delivery. Together they cover all Earth’s land surfaces, large islands, inland and coastal waters every five days at the equator. S-2A was launched June 2015 and S-2B March 2017. The thermal design has been analyzed and verified through thermal balance and vacuum tests to validate the thermal mathematical model. The paper describes the development of the mathematical model, the execution of the thermal tests in different facilities (IABG and ESTEC) and the corresponding correlation activities. The predicted and actual in-orbit performances are also reported. The thermal balance performed with the S-2A satellite showed already a good match of the thermal model to the as-built status. Nevertheless, some discrepancies were found on the Star Tracker Assembly and the Propulsion Plate. The thermal test of the S-2B satellite revealed further needs to improve the S-2A correlated thermal model. A summary of the most relevant parameters is given within this paper. The comparison of flight data during the in-orbit commissioning phase with the flight predictions performed with the correlated model show discrepancies in the order of only +/-3K for most units. The comparison of duty cycles shows deviations higher than 10% only for the units which already had higher discrepancies during the thermal balance correlation, namely the Star Tracker Assembly and the Propulsion Plate. The comparison of flight data shows a good similarity in temperature, stability and performance of both satellites. Continuous monitoring of the thermal behavior since launch is also revealing noticeable degradation of thermo-optical properties due to sun exposure, in particular on the Star Trackers. 1 Thermal Engineering & AIT, Thermal Engineering FHN, [email protected] 2 Thermal Analysis and Design Expert, Thermal Engineering FHN, [email protected] 3 System Architecture & IV&V, Mechanical Systems Germany, [email protected] Copyright © 2018 Airbus Defence and Space GmbH MMFU = Mass Memory and Formatting Unit Nomenclature MSI = Multi Spectral Instrument BOL = Beginning Of Life OBC = On-Board Computer CDR = Critical Design Review OCP = Optical Communication Payload CESS = Coarse Sun and Earth Sensor PCDU = Power Control and Distribution Unit CFRP = Carbon Fiber Reinforced Plastic PDR = Preliminary Design Review EOL = End Of Life PFM = Proto Flight Model ESA = European Space Agency PROP = Propulsion ESTEC = European Space Research and Technology RIU = Remote Interface Unit Centre S-2 = Sentinel-2 FM = Flight Model S/A = Solar Array FPA = Focal Plane Assembly S/C = Spacecraft GEU = Gyro Electronic Unit (IMU electronics) SADM = Solar Array Drive Mechanism GFU = Gyro-Fibre Unit (IMU sensor) SBT = GPS Satellite Binary Time GL = Linear Coupling STR = Star Tracker GMM = Geometrical Mathematical Model TB = Thermal Balance GPS = Global Positioning System TBTV = Thermal Balance and Thermal Vacuum GPSR = GPS Receiver TCS = Thermal Control System GR = Radiative Coupling TICD = Thermal Interface Control Document IABG = Industrieanlagen-Betriebsgesellschaft mbH TMM = Thermal Mathematical Model IMU = Inertial Measurement Unit TV = Thermal Vacuum IOC = In-Orbit Commissioning VCU = Video and Compression Unit IOV = In-Orbit Verification VDA = Vacuum Deposited Aluminium LCT = Laser Communication Terminal WSA = Weltraumsimulationsanlage (Space LEOP = Launch and Early Orbit Phase Simulation Chamber) LSS = Large Space Simulator XBS = X-Band Subsystem MLI = Multi Layer Insulation I. Introduction HIS chapter provides a comprehensive overview of the S-2A/B satellite design with focus on the platform side. T The satellite is characterized by a modular configuration enabling as much as possible parallel development and integration of its main subassemblies. The main structure of the platform is constituted by an aluminum frame structure with an intermediate floor, which divides the satellite in a lower electronics and an upper instrument compartment. The instrument compartment is closed by 4 access/radiator panels on Y&Z-sides. The radiator areas are optimized to the minimum possible sizes, reducing the need for operational heater power in cold case conditions. All external surfaces not used as radiators are thermally insulated to the maximum extent possible by means of Multi-Layer Insulation (MLI) blankets, limited by a number of unavoidable heat leaks, as there are the launcher I/Fs, thrusters, sensor apertures and antennas. An electrical heater system consisting of independently controllable heater lines is implemented to maintain minimum unit and payload temperatures during operating and non-operating phases as well as for temperature stabilization of the Star Tracker (STR) and Giro Fibre Unit (GFU) assemblies. Figure I-1: Satellite Design Overview. 2 International Conference on Environmental Systems The top floor, on +X-side, is providing accommodation for the Multi-Spectral Instrument (MSI), as shown in Figure I-1, which is iso-statically mounted by means of six CFRP struts on top of an intermediate CFRP panel which is again attached by 9 brackets to the platform top panel. The MSI electronic units are mounted close to the instrument on the rear side of the top panel. To achieve the required pointing accuracy, the alignment critical sensor assemblies carrying Star Trackers and IMU sensor (GFU) are accommodated directly on the MSI, as shown in Figure I-2. The Platform unit equipment compartment accommodates (by means of 4 sandwich panels and the intermediate floor) the electronic equipment. The satellite units are mounted on the inside of the platform aluminium sandwich panels with exception of the batteries which are mounted thermally insulated on the outside allowing for independent thermal control and late integration. The units have been distributed to equalize as much as possible the thermal dissipation, but also to keep the center of gravity balancing required by the launcher. The main dissipating units are the VCUs, PCDU, RIU and MMFU. The X-Band Subsystem as a whole also exhibits a substantial orbit average as well as peak heat dissipation. Radiators on Zenith and Deep Space side Radiators on Nadir and Sun side STR Radiators SBT & GEU Radiator VCU Radiator Access Panel Radiator Access Panel Radiator SADM Radiator LCT Radiator Battery Module Radiators OBC Radiator MMFU Radiator PCDU Radiator XBS Radiator GPSR Radiator WDE Radiator Figure I-2: Satellite External Thermal Model. A milled aluminium base plate is providing the interface to the cylindrical adapter with the clampband I/F connecting the satellite to the launch vehicle adapter. The propulsion subsystem is designed as a separate module which will be fully integrated prior to the integration into the platform. It is thermally decoupled from the satellite structure to the maximum extent possible. Since the minimum operational temperature of the RCS equipment is higher than the minimum operational temperature of the other internal units the propulsion module is covered with MLI to reduce radiative heat exchange and to minimize heat power for cold case. The pipework lies within an enclosure formed by the propulsion plate and the external MLI. Its temperature is radiatively controlled by means of heaters installed on the propulsion plate. The single wing solar array is designed such that its rotation axis is in line with both center of gravity of the S/C and the deployed wing, which minimizes the in-flight disturbances. It is considerably reducing the heat rejection capability of the PCDU and SBT/GEU radiators. II. TB/TV Test The Sentinel-2 satellites were subjected to thermal balance (TB) and thermal vacuum (TV) tests. The TB test is only performed on the S-2A PFM satellite as part of a combined TB/TV test while the S-2B FM satellite is just subjected to a TV test. However, in order to verify the performance and workmanship of the TCS, at least one FM TV test 3 International Conference on Environmental Systems phase had to be executed in exactly the same way as on the PFM. The two test campaigns were performed at IABG and ESTEC respectively. The S-2 satellite level thermal test objectives are summarized below: Thermal Balance Test Objectives: qualification of the platform thermal design & hardware w.r.t. performance and workmanship validation of the Thermal-Mathematical Model (TMM) of the Sentinel-2 platform by means of steady-state (thermal balance) as well as transient (orbit simulation) test phases verification of platform to payload / unit thermal interfaces confirmation of radiator and heater sizing already verified on instrument/equipment level Thermal Vacuum Test Objectives: verification of the functional performance of the satellite under extreme in-orbit environmental conditions by means of Abbreviated
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